Adjustable blade-to-hub lead-lag damper attachment

ABSTRACT

An aircraft includes a rotor blade and a rotor hub system. The rotor hub system includes a lead-lag damper having a rod end and being operably associated with the rotor blade; a blade adapter having a first arm and an opposing second arm; a pin carried by the blade adapter and configured to secure in position via the first arm and the second arm; and an actuator secured to the pin and configured to adjust the position of the lead-lag rod end relative to the first arm and the second arm.

BACKGROUND

1. Field of the Invention

The present application relates generally to dampers, and morespecifically, to an adjustable lead-lag damper.

2. Description of Related Art

In articulated rotors, blade lead-lag, feathering and flapping motionswill all potentially contribute to one-per-revolution damper motion,which affects hub and blade loads as well as damping of the lead-lagmode.

Helicopter with articulated rotors require the use of lead-lag damper toremain stable on the ground and in the air. However, in addition toproviding damping to the rotor, lead-lag dampers will also generateloads, which are generally a function of the one-per-revolution dampermotion (more motion, more load). Furthermore, it is known thatone-per-revolution damper motion in conjunction with motion at the lagmode frequency, reduces the lead-lag damping of the rotor system (moremotion, less damping). Therefore, minimizing one-per-revolution dampermotion would be beneficial in terms of both reducing loads andincreasing the dual frequency damping of the lead-lag mode.

This one-per-revolution damper displacement is essentially related tothe rotor geometry (location of the blade pivot points and damperattachment points) and the blade lead-lag, feathering and flappingmotions. In addition, the phase shift between these blade motions isimportant, as different phasing could result in either an additive orcancelling effect, increasing or reducing overall damper displacement.

Although the foregoing developments in the field of lead-lag dampersrepresent great strides, many shortcomings remain.

DESCRIPTION OF THE DRAWINGS

The novel features believed characteristic of the embodiments of thepresent application are set forth in the appended claims. However, theembodiments themselves, as well as a preferred mode of use, and furtherobjectives and advantages thereof, will best be understood by referenceto the following detailed description when read in conjunction with theaccompanying drawings, wherein:

FIG. 1 is a side view of a helicopter according to a preferredembodiment of the present application;

FIGS. 2-5 are perspective views of a rotor hub system in accordance witha preferred embodiment of the present application;

FIGS. 6A and 6B are perspective views of a rotor hub system inaccordance with an alternative embodiment of the present application;and

FIG. 7 is a perspective view of a rotor hub system in accordance with analternative embodiment of the present application.

While the system and method of the present application is susceptible tovarious modifications and alternative forms, specific embodimentsthereof have been shown by way of example in the drawings and are hereindescribed in detail. It should be understood, however, that thedescription herein of specific embodiments is not intended to limit theinvention to the particular embodiment disclosed, but on the contrary,the intention is to cover all modifications, equivalents, andalternatives falling within the spirit and scope of the process of thepresent application as defined by the appended claims.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

Illustrative embodiments of the apparatus and method are provided below.It will of course be appreciated that in the development of any actualembodiment, numerous implementation-specific decisions will be made toachieve the developer's specific goals, such as compliance withsystem-related and business-related constraints, which will vary fromone implementation to another. Moreover, it will be appreciated thatsuch a development effort might be complex and time-consuming, but wouldnevertheless be a routine undertaking for those of ordinary skill in theart having the benefit of this disclosure.

The system of the present application overcomes the abovementionedproblems commonly associated with conventional articulated rotor hubassemblies. The system discussed herein provides a means of changing thephasing of the different contributions to damper displacement tominimize damper one-per-revolution motion. This is accomplished throughan adjustable damper attachment to the hub and/or blade. Theadjustability of the damper attachment points is what would allow thereduction in loads and the increase in dual frequency lead-lag damping.Such adjustability is provided by appropriate hardware added to theattachment locations. The adjustment could be purely mechanical,performed let's say before flight during the flight test program toidentify the optimal configuration, or it could be part of an activesystem that allows adjustments to be performed in-flight. Furtherdetailed description of these features are provided below andillustrated in the accompanying drawings.

The system and method of the present application will be understood,both as to its structure and operation, from the accompanying drawings,taken in conjunction with the accompanying description. Severalembodiments of the system are presented herein. It should be understoodthat various components, parts, and features of the differentembodiments may be combined together and/or interchanged with oneanother, all of which are within the scope of the present application,even though not all variations and particular embodiments are shown inthe drawings. It should also be understood that the mixing and matchingof features, elements, and/or functions between various embodiments isexpressly contemplated herein so that one of ordinary skill in the artwould appreciate from this disclosure that the features, elements,and/or functions of one embodiment may be incorporated into anotherembodiment as appropriate, unless described otherwise.

Referring now to the drawings wherein like reference characters identifycorresponding or similar elements throughout the several views, FIG. 1depicts an aircraft 101 in accordance with a preferred embodiment of thepresent application. In the exemplary embodiment, aircraft 101 is ahelicopter having a fuselage 103 and a rotor system 105 carried thereon.A plurality of rotor blades 107 is operably associated with rotor system105 for creating flight. In the exemplary embodiment, only two blades107 are shown; however, it will be appreciated the features of thesystem discussed herein are preferably operably associated with three ormore blades with an articulated soft-in-plane rotor hub assembly.

Although shown associated with a helicopter, it will be appreciated thatthe system of the present application could also be utilized withdifferent types of rotary aircraft having articulated soft-in-planerotor hub assemblies.

FIGS. 2-5 depict various view of a rotor hub system 301 in accordancewith a preferred embodiment of the present application. In thecontemplated embodiment, system 301 overcomes one or more of theproblems commonly associated with conventional articulated rotor hubassemblies. To achieve these features, system 301 includes a lead-lagdamper 303 that is adjustable in height relative to a damper attachmentjoint 305. The selective placement of the lead-lag damper 303 relativeto damper attachment joint 305 overcomes the problems discussed abovewith respect to conventional articulated rotor hub assemblies.

System 301 includes a blade adapter 307 having a first arm 309 and asecond arm 311 configured to receive a pin 316, which in turn secures toa rod end 315. Rod end 315 is retained in position via a first spacer313 and a second spacer 314 configured to receive the pin 316therethrough. Thus, in the exemplary embodiment, arms 309, 311 alongwith pin 316 create damper attachment joint 305 that the lead-lag damper303 is secured thereto via a spherical-bearing rod end 315.

One of the unique features believed characteristic of the presentapplication is the selective adjustment of rod end 315 relative to arms309, 311, as depicted with direction arrow D1. The gap 317 between firstarm 309 and rod end 315 is represented by a distance D2, while the gap319 between second arm 311 and rod end 315 is represented by a distanceD3. In the exemplary embodiment, the distances D2 and D3 are equal toeach other, thereby pacing the rod end at a neutral position. Movementof the rod end 315 changes the position to either positive or negative,as will be shown and discussed below.

Referring specifically to FIGS. 4 and 5, system 301 is further providedwith a spacer 501 configured to receive pin 316 and sit between the arms309, 311 of the blade adapter and the rod end 315. In one contemplatedembodiment, spacer 501 could include an aperture (not shown) extendingthrough the longitudinal length of the spacer and configured to receivethe pin 316 therethrough. Alternative embodiments could includes aspacer that is rigidly attached to the blade adapter 307 and/orremovably secured to pin 316 for selective adjustment after use.

It will be appreciated that the spacer 501 is configured to retain rodend 315 in an offset position relative to arms 309, 311. Thus, in theexemplary embodiments of FIGS. 4 and 5, distances D2 and D3 are nolonger equal to each other and the rod end 315 is considered to be in anoffset position. In FIG. 4, the rod end 315 is considered in a negativeoffset position where distance D2 is greater than distance D3, whileFIG. 5 shows rod end 315 in a positive offset position wherein distanceD2 is less than distance D3.

It will be appreciated that the distances D2, D3 relative to each otheris selectively tailored to achieve optimal flight performance. In thecontemplated method of use, spacer longitudinal length is determinedprior to flight testing. Thus, the spacer 501 does not allow forreal-time adjustment of D2, D3 during flight.

Another unique feature believed characteristic of the presentapplication is the ability to selectively adjust distances D2, D3 whilethe aircraft is in flight. To achieve this feature, a system 701includes an actuator 703 operably associated with a pin 602 andconfigured to move pin 602 in direction D1, as depicted in FIGS. 6A and6B. The movement of pin 602 allows selective adjustment of D2 and D3during flight. The control of actuator 703 can be achieved autonomouslywith a flight control computer (not shown) and/or manually. In thecontemplated embodiment, pin 602 is a helical pin having threads formoving the rod end in direction D1. However, alternative embodimentscould utilizes a telescoping pin and/or other suitable means to move rodend 315 in direction D1.

In FIG. 7, a perspective view of a system 801 is shown in accordancewith an alternative embodiment of the present application. System 801 issubstantially similar in function to system 301. However, system 801 isconfigured such that the arms 309, 311 are positioned such that no gapis formed between the arms and the rod end 315. To achieve this feature,arm 309 and/or arm 311 is adjusted relative to the blade adapter. Thus,the surface 803 of blade adapter 307 is spaced apart from surface 805 offirst arm 309 at a distance D4 so as to create equal distances D2, D3between the arms and the rod end.

It is apparent that a system and method with significant advantages hasbeen described and illustrated. The particular embodiments disclosedabove are illustrative only, as the embodiments may be modified andpracticed in different but equivalent manners apparent to those skilledin the art having the benefit of the teachings herein. It is thereforeevident that the particular embodiments disclosed above may be alteredor modified, and all such variations are considered within the scope andspirit of the application. Accordingly, the protection sought herein isas set forth in the description. Although the present embodiments areshown above, they are not limited to just these embodiments, but areamenable to various changes and modifications without departing from thespirit thereof.

What is claimed is:
 1. A rotor hub system for an aircraft, comprising: alead-lag damper operably associated with a rotor blade of the aircraft;a blade adapter, having: a first arm and an opposing second arm; a pincarried by the blade adapter and secured in position via the first armand the second arm, the pin supporting a rod end of the lead-lag damper;and a first interchangeable spacer and a second interchangeable spacer,both spacers being configured to receive the pin therethrough andconfigured to retain the rod end of the lead-lag damper secured to thepin at a selected distance relative to the first arm and second arm;wherein the pin is a telescoping pin; and wherein translation of thetelescoping pin causes translation of the rod end between the first armand the second arm.
 2. The system of claim 1, wherein the first spacerand the second spacer each have a longitudinal length; and wherein thelongitudinal length of the first spacer is less than the longitudinallength of the second spacer, thus retaining the rod end at a non-equaldistance between the first arm and the second arm.
 3. The system ofclaim 1, wherein the first spacer and the second spacer each have alongitudinal length; and wherein the longitudinal length of the firstspacer is equal to the longitudinal length of the second spacer, thusretaining the rod end at an equal distance between the first arm and thesecond arm.
 4. The system of claim 1, wherein the rotor hub system is anarticulate rotor hub system.
 5. The system of claim 4, wherein thearticulated rotor hub system is a soft-in-plane assembly.
 6. The systemof claim 1, wherein the rotor hub system is operably associated with atleast three rotor blades.
 7. An aircraft, comprising: a rotor blade; anda rotor hub system, having: a lead-lag damper having a rod end and beingoperably associated with the rotor blade; a blade adapter, having: afirst arm and an opposing second arm; a pin carried by the blade adapterand secured in position via the first arm and the second arm, the pinsupporting a rod end of the lead-lag damper; an actuator secured to thepin and configured to adjust the position of the lead-lag rod endrelative to the first arm and the second arm; and a first spacer and asecond spacer, both spacers being configured to receive the pintherethrough and configured to retain the rod end of the lead-lag dampersecured to the pin at a selected distance relative to the first arm andsecond arm.
 8. The aircraft of claim 7, wherein the aircraft is ahelicopter.
 9. The aircraft of claim 7, wherein the rotor hub system isan articulate rotor hub system.
 10. The aircraft of claim 9, wherein thearticulated rotor hub system is a soft-in-plane assembly.
 11. Theaircraft of claim 7, wherein the rotor hub system is operably associatedwith at least three rotor blades.
 12. The aircraft of claim 7, whereinthe actuator is autonomously controlled.
 13. A method to increase flightperformance of an aircraft, comprising: providing the aircraft of claim7; and adjusting the position of the rod end relative to the first armand the second arm via an actuator.
 14. The method of claim 13, furthercomprising; autonomously controlling the adjustment of the rod end. 15.The method of claim 13, wherein the pin is a threaded helical pin;wherein the process of adjusting the position of the rod end is achievedvia the threaded helical pin operably associated with the rod end of theactuator.
 16. The method of claim 15, wherein the actuator rotates thehelical threaded rod.
 17. The method of claim 13, wherein the pin is atelescoping pin; wherein the process of adjusting the position of therod end is achieved via the telescoping pin operably associated with therod end and the actuator.
 18. The method of claim 17, wherein theactuator translates the telescoping pin.